The field of the present disclosure relates generally to turbine engines and, more particularly, to combustor assemblies for use in turbine engines.
In a gas turbine engine, air pressurized in a compressor is mixed with fuel in a combustor to generate hot combustion gases. Energy is initially extracted from the gases in a high pressure turbine (HPT) that powers the compressor, and subsequently in a low pressure turbine (LPT) that powers a fan in a turbofan aircraft engine application, or powers a propeller via a gearbox in a turboprop aircraft application, or powers a rotor via a gearbox in a turboshaft helicopter application, or powers an external shaft for marine and/or industrial applications. Generally, engine efficiency increases as the temperature of combustion gases is increased. However, the increased gas temperature increases the operating temperature of various components along the gas flowpath, which in turn increases the need for cooling such components to facilitate extending their useful life.
For example, known combustors include an annular combustion liner that requires cooling during operation of the gas turbine engine. Furthermore, known reverse flow combustors used in gas turbine engines generally include a large surface area that requires additional cooling. At least some gas turbine engines use compact axial through flow combustors to reduce the amount of surface area requiring cooling. However, at least some known gas turbine engines have a limited amount of axial space between the compressor and the HPT. In such configurations, axial through flow combustors may not fit in the allotted space, thus requiring the use of a reverse flow combustor having a reduced surface area and sufficient cooling features.